The present invention relates generally to gas turbine engines, and, more specifically, to turbine cooling therein.
In a gas turbine engine, air is pressurized in a compressor, mixed with fuel in a combustor and ignited for generating hot combustion gases, which flow downstream through one or more turbine stages for extracting energy therefrom. A high pressure turbine (HPT) firstly extracts energy from the gases for powering the compressor. And, additional energy is typically extracted from the gases by a low pressure turbine (LPT) which typically powers a fan disposed upstream from the compressor.
The HPT includes a stationary turbine nozzle which directly receives the combustion gases from the combustor for redirecting the gases into a row of rotary turbine blades extending radially outwardly from a rotor disk. The nozzle includes a plurality of circumferentially spaced apart stator vanes which complement the performance of the rotor blades.
Both the vanes and blades are suitably configured as a airfoils which cooperate for maximizing efficiency of extraction of energy from the combustion gases which flow thereover. The vane and blade airfoils have generally concave pressure sides and opposite, generally convex suction sides which extend axially between corresponding leading and trailing edges thereof and radially over their radial span.
The nozzle vanes extend radially between annular outer and inner bands which confine the combustion gases therebetween. The blade airfoils extend from their radially inner roots to their radially outer tips which are spaced closely radially inwardly from a surrounding annular turbine shroud. The shroud is stationary and defines the outer boundary for the combustion gases which flow past the rotating blade airfoils.
Since the stator vanes, rotor blades, and turbine shrouds are directly exposed to the combustion gases, they require suitable cooling for maintaining their strength and ensuring suitable useful lives thereof. These components are typically cooled by channeling thereto corresponding portions of air bled from the compressor which is substantially cooler than the hot combustion gases. Various cooling techniques are used in cooling gas turbine engine components. Film cooling is one technique wherein air is channeled through inclined film cooling holes to form a film of cooling air between the outer or exposed surfaces of the components and the hot combustion gases which flow thereover.
Impingement cooling is another technique wherein the cooling air is initially directed substantially normal to the inner surfaces of these components in impingement thereagainst for removing heat therefrom by convection heat transfer. The inner surfaces may be smooth for impingement cooling, or may include three dimensional turbulators in the form of cylindrical pins, bumps, or dimple depressions. These turbulators increase the effective surface area of the inner surfaces from which heat may be extracted. The turbulators are typically small in size for reducing any adverse pressure drop caused thereby for ensuring cooling efficiency.
Since turbine vanes, blades, and shrouds are formed of high strength metals, they are typically manufactured by casting for achieving maximum material strength and precision of the small features thereof, including any turbulators which may be used therein.
The vanes and blades are hollow for channeling therethrough the cooling air in several radially extending passages. The passages may be individually fed with cooling air or may be arranged in serpentine legs through which the cooling air flows. Impingement cooling for the vanes is typically provided by placing perforated impingement baffles inside corresponding internal passages therein. The cooling air is first channeled inside the baffle and then laterally through its perforations for impingement against the inner surface of the vane.
Since turbine blades rotate during operation, an integral rib or bridge may be provided between its pressure and suction sides for defining an integral baffle having holes or perforations through which the cooling air is directed in impingement against the inner surface of the blade airfoil, typically along the leading edge.
Both the vane and blade airfoils may be similarly cast in view of their common airfoil configurations with internal radial passages. The internal passages are defined by corresponding ceramic cores surrounded by wax which defines the configuration of the final airfoil. The wax is then surrounded by a ceramic shell, and subsequently removed in the lost wax method. Molten metal is then poured between the shell and core and solidifies in the form of the desired airfoil. The ceramic shell and cores are then removed to expose the cast airfoil.
The ceramic cores themselves are produced in a separate casting process using a metallic core die precisely formed with the mirror features to be produced in the outer surface of the core. A typical core die may be formed in two or more halves with an internal passage being defined therebetween and extending along the span axis thereof. A ceramic slurry or paste is injected under significant pressure in the open end of the die to fill the die, after which the resulting ceramic core is removed and cured.
The same core die is used repeatedly for casting multiple copies of the airfoils. However, the injection of the ceramic slurry into the die eventually leads to wear therein. Wear is most pronounced for three dimensional features such as the turbulators for enhancing impingement cooling, which turbulators of the core die are abraded over extended use. Once the die is worn, a new die must be manufactured at considerable expense.
Accordingly, it is desired to provide improved impingement cooling features in a turbine component, which can reduce core die wear in a preferred embodiment.